1. Field of the Invention
The present invention relates generally to apparatus for simulating, as if airborne and exposed to the air stream, aerodynamic heating and cooling effects on equipment intended for placement externally of an aircraft or missile.
2. Description of the Prior Art
Apparatus used to simulate aircraft or missile operating conditions while on the ground have long been used. A number of these which have been disclosed in the patent literature will now be discussed.
U.S. Pat. No. 3,960,000 to Barnett et al., "Flight Simulator for Missiles", includes means for directing heat toward the nose cone of a missile while imparting rotation to the missile, simulating the flight conditions of spinning missiles. The heat source typically takes the form of an oxygen gas flame controlled by a controller. An IR detector cell monitors the actual heat (temperature) reached at the surface of the nose cone.
The present invention differs from Barnett et al. in several key respects. The goal of Barnett et al. is to heat the surface of the nose cone to temperatures predicted or measured during test. It is not a goal to also match the correct corresponding convective heat transfer coefficient with the flight conditions. This is not critical in the typical fiberglass or other insulative type nose cone tipped missile where the heat from Internal electronics would not be rejected through the nose cone. Instead, one of the nose cone's main purpose is to protect by insulation the electronics (a fuse in the case of Barnett et al.) from the aerodynamic heating experienced by a high speed missile.
In contrast, the present invention has, as one object, the goal of accurately producing the convective heat transfer coefficient produced by the air stream in cases where the air stream is used to cool the electronic equipment which has a surface or surfaces exposed to the air stream for this purpose. In less technical terms, correctly simulating this coefficient will correctly model the resistance of heat entering from air stream (aerodynamic heating) or leaving the flight vehicle into the air stream (aerodynamic cooling) when internal heat is being generated. In practical cases where the heat is being rejected into the air stream i.e. from electronics, the surface of a flight vehicle will be heated by both the electronic heat sources and the air stream until the surface temperature is greater than the air stream adiabatic wall temperature so that the generated heat is rejected at the rate of the electronic dissipation.
Considering the effects of aerodynamic heating and the inherent temperature limitations of electronics, this skin temperature is limited to a maximum of about 100.degree. C. Corresponding flight speeds would be limited to high subsonic at low altitudes to moderate supersonic speeds of Mach=2.0 at altitude. This of course is the speed range for the majority of all commercial and military aircraft and one object of the present invention is provide the aero-heating and cooling effects for equipment for this broad spectrum of flight vehicles.
In contrast, by the nature of its oxygen gas heat source, Barnett et al. addresses a smaller class of applications than the present invention, primarily high speed missiles reaching Mach=2.5 (adiabatic wall temperatures of 220.degree. C. or higher) where the high temperatures provided by this type of heat source are required.
U.S. Pat. No. 3,709,026 to Rhodes et al., "Apparatus and Method for Simulating Spacecraft Erosion" discloses missile/spacecraft testing equipment. This invention has the means to simulate the effects of high speed flow, including heating and impact with particulate in the atmosphere on a test object made of materials which are under consideration for the external surfaces of a space craft entering or leaving the earth's atmosphere. Particulate include dust, rain, ice particles, vapor and micro-meteorites. Effects of the complete flight environment on heat shield materials, antennas, optical windows and like materials which are of interest to designers of high speed rocket vehicles entering or leaving the earth's atmosphere can be simulated.
Specifically the Rhodes et al. invention has a plurality of accelerator nozzles means positioned on and protruding through the walls of spherical thick walled housing made of metals which is capable of withstanding the test pressures and high temperatures (up to 5,000.degree. F.). Each of nozzle accelerators can independently accelerate air under high pressure to stream of high velocity air heated to high temperatures, simulating atmospheric heating, while injecting any of the above particulate as desired into the stream to impact on the test object.
The test object is mounted to a stage with associated mechanisms which can position the test object sequentially in front of a group of these accelerator nozzles, each operating at various conditions to simulate the time history of the trajectory of the flight vehicle entering or Leaving the earth's atmospheric conditions to simulate the time history of the trajectory of the flight vehicle entering or leaving the earth's atmosphere.
The present invention differs substantially from Rhodes et al. in function. The air stream emanating from the accelerating nozzles in the patent immerses the test object in the full free stream air velocity that would be encountered in flight. To accomplish this necessitates that the nozzle and emanating air stream is slightly larger than the test object. Since the nozzle is only a fraction of the size of the test chamber for flow purposes only a small scale model could be tested with a reasonably sized bench top test chamber. However, one goal of the present invention is to be able to test full scale actual electronic equipment in aircraft and missiles where a portion or all of the equipment boundaries are exposed to the air stream. If the structure of Rhodes et al. is reduced to one large nozzle within the chamber, the flow rates required will be at least an order of magnitude larger than the present invention, functioning like a wind tunnel requiring use of much larger compressors, greater power consumption and a much larger and expensive chamber to produce the actual flight conditions, than required by the present invention. Additionally the present invention can simulate changing flight conditions without use of the expensive positioning mechanisms used in Rhodes et al. This is accomplished in the present invention by changing the pressure and heating of the air flow supplied to the test article housing.
U.S. Pat. No. 3,230,764 to Bloxsom, Jr. Et al., "Method of Determining Heat Transfer Rates and Temperature Conditions", is directed toward an optical sensor means for measuring the heat transfer rates and temperatures reached in the very short test time available for test models in a hypersonic pulse type wind tunnel. This type of wind tunnel can only provide test simulation on the order of a millisecond. The optical sensor means is based on the change in the refractive index of many plastic compositions including acrylics with temperature. The test model is coated with this type of sensor material and post test examination can determine the amount of heating by the depth of the plastic experiencing a change of refractive index. Optical interference methods are used for this determination. As opposed to the present invention, Bloxsom et al. does not simulate the heat transfer between the air stream and test model but instead measures the result and in this manner is distinguished from the present invention.
U.S. Pat. No. 3,121,329 to Bennett, "Simulation of Reentry Conditions". employs means which produces streams of partially ionized gas at speeds characteristic of orbiting satellites within an evacuated space simulating enviromnental conditions of the upper atmosphere and outer space. These means include projecting a source of positive nitrogen ions through the cover into the cylindrical member at about 5000 volts into a nitrogen gas which is then accelerated to orbital velocities by collisions with the ions. A nozzle then directs this stream of accelerated nitrogen toward the front of the test model whereby reentry gas dynamic conditions can be closely simulated.
The Bennett invention is a very complex and expensive apparatus which uses a combination of expedients including use of using high voltage means to produce and accelerate nitrogen ions and thereby gaseous nitrogen to at least orbital velocities. The present invention has as a goal and the means to simulate aerodynamic heating and cooling at the velocities of the majority of air flight vehicles rather than space vehicles i.e. subsonic through low supersonic rather than reentry and in this broad manner may be distinguished from Bennett.
U.S. Pat. No. 3,027,760 to Holderer, "Adjustable Porous Walls for Wind Tunnels", has primary application to transonic wind tunnels and introduces variable perforation means in wall means to absorb shock waves emanating from a test model. A solid wall would reflect the shock waves back toward the model interfering with the intent to simulate the flow over the model that would occur in flight through the atmosphere. The perforations allow a means of escape of the compressed air from a shock wave into an outer chamber which absorbs the energy from the wave. The outer chamber requires significant air flow which is thereby minimized by the ability to vary the perforation size depending on the strength of the shock waves.
The invention of Holderer can be easily be distinguished from the present invention, in that the patent has an object to enhance the accuracy of the air flow in a wind tunnel. In contrast, the present invention has as an object to replicate the cooling and heating effects of air flow around a unit under test by the disclosed means without using a wind tunnel Again this approach will only use a small fraction of the air flow and compressor size and power of a wind tunnel.
U.S. Pat. No. 2,510,952 to Brewster, "Temperature Testing Chamber", discloses means to closely regulate the air temperature in a ground pressure testing chamber for equipment preventing sudden temperature excursions. The invention uses valve and ducting means to separately re-circulate from the chamber through the cooling and heating means where the temperature of the test chamber can be accurately raised or lowered without significant temperature overshoot. Brewster can be distinguished from the present invention in that its capability is limited to temperature control only and can not simulate the heating or cooling effects on a test article from the air stream during flight.
U.S. Pat. No. 2,439,806 to Heineman, "Testing Chamber", discloses means to perform testing of equipment and personnel at simulated high altitude conditions. The invention of Heineman introduces a thin metallic membrane into the insulated walls to prevent migration of moisture and damage to the insulation. The Invention of Heineman can be distinguished from the present invention in that its capability is limited to temperature, altitude, and humidity control and can not simulate the heating or cooling effects on a test article from the air stream during flight.
U.S. Pat. No. 2,309,938 to Diserens et al., "Cooling System for Wind Tunnels or Similar Enclosures", discloses means to cool large air flow rates required by wind tunnels down to simulated high altitude conditions without adding significant blockage and flow resistance in the form of cooling coils or other direct contact cooling apparatus in the wind tunnel loop. The means includes an external air liquefaction system which accumulates liquid air and then sprays the liquid at as high rates as needed through a distribution manifold into the air upstream of the test section. Compared to cooling coils, the spray system adds little flow impedance, while delivering a substantial cooling effect.
Diserens et al., can be distinguished from the present invention in that the means and goals of the patent are limited to providing cooling of air flowing through the wind tunnel In contrast, the present invention simulates the cooling and heating effects of the air flow on airborne equipment at high speeds. In addition, the air flow rate needed by the present invention is only a small fraction of that required by a wind tunnel.
It was with knowledge of the foregoing state of the technology that the present invention has been conceived and is now reduced to practice.